In order to allow for the use of the various mission instruments of a satellite, such as telecommunication systems, it should be possible to control the position and the orientation of the satellite in orbit. For this, control systems are implemented for maintaining, on the one hand, the orientation of the satellite relative to the Earth, that is to say the attitude control, and, on the other hand, its position in orbit relative to a desired ideal position, that is to say orbit control. For example, in the case of geostationary satellites, orbit control seeks to limit the inclination relative to the equatorial plane, to limit the eccentricity of the orbit, and to limit the drift in the longitudinal position of the satellite relative to Earth. For this, thrusters are positioned at various points on the satellite to correct the trajectory at more or less short intervals by applying a force to the satellite. These station-keeping operations require the availability of a sufficient reserve of fuel throughout the life of the satellite.
A satellite is placed in orbit by the combination of a space launch vehicle and its own propulsion systems. The launch vehicle transports and releases the satellite on a first Earth orbit called transfer orbit, the perigee of which is generally low; once in this first orbit, a propulsion system of the satellite takes over to transport the satellite to its final orbit. Generally, this transfer is performed by means of a main satellite thruster PSP that consumes a chemical fuel of ergol or propergol type, delivering a thrust of high power making it possible to rapidly reach the final orbit.
Once in station, a number of lower power secondary thrusters ensure that the satellite is kept in position in the orbit. For this, ergol chemical thrusters or electrical thrusters can be used. In an electrical thruster, of plasma thruster or ion thruster type, xenon atoms are ionized by collision with electrons, creating xenon ions. The thrust is generated when the charged xenon ions are accelerated from the thruster by an electromagnetic field. Although costly and of great initial weight, the effectiveness of the thruster, or its capacity to generate force by mass ejection, also called specific impulsion, is substantially greater than that of chemical thrusters.
In the known systems chemical thrusters and electrical thrusters are positioned at various points on the structure of the satellite to address all the needs of the mission, from transport from the transfer orbit to keeping in orbit throughout the life of the satellite. The propulsion systems that are thus implemented have the drawback of a high cost and high weight, different thrusters and fuel. These drawbacks limit the payload capacity of the satellite.
According to the known prior art, an orbit control system seeks to control the position of the satellite through six orbital parameters. FIG. 1 represents a geostationary satellite 10 in orbit 11 around the Earth 12. The orbit 11 is inclined by an angle θ relative to the equatorial plane 13 which contains the ideal geostationary orbit 14. The orbit 11 of the satellite intersects the equatorial plane 13 at two points 15 and 16, commonly called orbital nodes. The six orbital parameters used to qualify the position of a satellite are also known: the half major axis, the eccentricity, the inclination, the argument of the ascending node, the argument of the perigee, and the true anomaly. The orbit control consists in quantifying these orbital parameters and performing the necessary operations by means of the onboard propulsion systems, to keep the satellite in a predefined zone around an ideal position. As an example, for a geostationary satellite, a drift window of plus or minus 0.1°, representing a width of almost 150 km, is allocated around a target position.
A routine architecture of a satellite 10, as represented in FIG. 2, comprises a parallelepipedal structure 20 onto which are fastened various devices useful to the piloting of the satellite 10 and to its mission. Telecommunication instruments 21 are installed on a face 22 whose orientation is kept facing the Earth, commonly called Earth face. On an opposite face 23, commonly called anti-Earth face, the main satellite thruster PSP is positioned, notably ensuring the thrust needed for the transfer from low orbit to the final orbit. On two opposing lateral faces 24 and 25, commonly called north face and south face, because of their orientation relative to the equatorial plane, there are positioned two sets of solar panels 26 and 27 making it possible to supply electrical energy to the onboard systems. Various devices can be installed on the lateral faces 28 and 29, commonly called east face and west face for their orientation relative to a Earth longitude. Maintaining a constant orientation of the satellite relative to the Earth is necessary to the successful completion of the mission of the satellite, for example for the orientation of the solar panels 26 and 27 or the pointing of the telecommunication systems 21 towards the Earth. This is done by means of an attitude control system. A number of attitude control systems are known that are suitable for detecting and correcting the orientation errors. Thus, the orientation of the satellite can be measured by means of a sensor assembly, comprising, for example, a sensor directed towards the Earth, positioned on the Earth face, for a measurement on two axes pitch and roll, relative to the Earth and a set 30 of gyroscopes to detect speeds of rotation on three axes. From these measurements, satellite orientation corrections about the center of gravity of the satellite can be made, for example by means of a set of inertia wheels 31 or gyroscopic actuators.
A satellite equipped with such a system allowing for attitude control is said to be stabilized on three axes. Typically, by controlling the speed of rotation and the orientation of the inertia wheels, it is possible to correct an orientation error within a reference trihedron linked to the satellite. Hereinbelow, Z designates an axis directed towards the Earth, also called yaw axis, Y an axis at right angles to the orbit and oriented in the direction opposite to the kinetic moment of the orbit (towards the south for a geostationary orbit), also called pitch axis, and X an axis forming, with Y and Z, a direct orthogonal reference frame, also called roll axis, which is oriented according to the speed in the case of circular orbits.
For the orbit control, a number of thrusters are arranged on the structure 20 of the satellite 10. A first high power thruster PSP, making it possible to ensure the transfer between the initial Earth orbit (after release from the launch vehicle) and the final orbit, is positioned on the anti-Earth face 23. According to a known prior art, a first set of thrusters, comprising, for example, two thrusters 32 and 33 positioned on the north face and on the south face in proximity to the anti-Earth face, is used to control the inclination. A second set of thrusters, such as, for example, the thrusters 34 and 35, positioned on the east and west faces, is used to control the eccentricity and the drift. It is also known that controlling the inclination takes around five to ten times more fuel than controlling the eccentricity and drift. For this reason, the inclination control is generally performed by means of a plasma thruster, which consumes less fuel, whereas the thrusters dedicated to controlling the eccentricity and drift are more often than not chemical ergol thrusters.
As an example, a regular satellite with a dry weight of 2500 kg and accommodating a payload of 900 kg comprises a main thruster, two plasma thrusters for the inclination and eccentricity, and four ergol thrusters for the eccentricity and drift. Typically, 1700 kg of ergol are needed for the initial orbit transfer, and 220 kg of xenon are needed to ensure the orbit control of the satellite for a mission duration of approximately 15 years. Thus, the cost and the weight of the current propulsion systems limit the capacity to have a high payload onboard. It can also be noted that, in most known propulsion systems for controlling orbit, the various onboard thrusters in reality comprise two propulsive engines positioned side by side, for mission safety and reliability reasons. This redundancy, well known to those skilled in the art, is not represented in the figures but it is considered hereinbelow that a thruster may consist of one or more propulsive engines forming a propulsive set, and of which the thrust which can be delivered is identical, in orientation or in intensity.
FIGS. 3a, 3b and 3c illustrate the principle of orbit control for a satellite according to the known prior art. The structure 20 of the satellite 10 is represented by a side view, the east face being visible. The thruster 32 is linked to the north face of the structure 20 by means of a two-axis mechanism 40. The two-axis mechanism 40 allows the rotation of the thruster 32 relative to the structure 20 on a first axis parallel to the axis Y and a second axis parallel to the axis X. In FIGS. 3a to 3c, the two-axis mechanism 40 is a universal joint link produced by means of a first pivot link 41 of axis parallel to the axis Y and a second pivot link 42 of axis parallel to the axis X. The centre of mass of the satellite, situated inside the parallelepipedal structure 20, is referenced CM.
In FIG. 3a, the orientation of the thruster 32 makes it possible to direct the thrust of the thruster towards the centre of mass CM of the satellite. To perform an inclination correction manoeuvre, a technique known to those skilled in the art consists in firing the thruster 32 a first time in proximity to an orbital node, for example 15, then the thruster on the opposite side a second time in proximity to the opposite orbital node, 16 in the example. Thus, the thrust from the first firing of the thruster 32, oriented towards the centre of mass CM, displaces the satellite in a direction that has a Z component and a Y component. Twelve hours later, the thrust of the second firing at the opposite orbital node displaces the satellite in a direction having a Z component opposite the first firing, and which compensates the undesired effect on the eccentricity and a Y component, also opposite, but the desired inclination effects of which accumulate. Thus, two firings of equal intensity are performed at twelve hour intervals in proximity to the orbital nodes 15 and 16 make it possible to cancel the effect of the radial component to retain only a north-south correction. This known procedure allows for a daily correction of the inclination.
By this same technique, it is also possible, by applying a second thrust of intensity different from the first, to apply eccentricity corrections on an axis at right angles to the line joining the two orbital nodes 15 and 16. Techniques have also been developed to allow for eccentricity corrections on a second axis, by offsetting the firing of the thruster relative to the orbital node, but at the cost of less effective control of the inclination. To sum up, the known systems make it possible, by means of two thruster systems 32 and 33, to ensure the control of the inclination and the control of the eccentricity on an axis with no loss of optimization of the inclination control, or to ensure the control of the inclination and the control of the eccentricity on two axes with a loss of optimization of the inclination control. The drift control can be handled by these two thrusters. A regular satellite comprises, for this, four chemical ergol nozzles positioned on the east and west faces of the satellite.
The thruster systems 32 and 33 are also useful for managing the quantity of movement of the attitude control systems, as illustrated in FIGS. 3b and 3c. By applying a thrust off the centre of mass CM—in a plane Y-Z in FIG. 3b and off the plane Y-Z in FIG. 3c, a rotation torque is generated on the satellite—a role torque in FIG. 3b and a pitch and yaw torque in FIG. 3c. These two torques can be used to load or unload the inertia wheels along two axes. For example, when the speed of rotation of an inertia wheel reaches it limit speed, effort will be made to orient the thrust deliberately off the centre of mass CM so as to generate, in addition to the desired displacement of the satellite, a torque making it possible to desaturate the inertia wheel, or more generally, the problem will be anticipated by reducing the kinetic moment to desired values at the time of each manoeuvre. These desired values obviously possibly being zero, but also a value shrewdly defined in such a way as to anticipate the trend of the kinetic moment between two manoeuvres under the effect notably of the solar radiation pressure.
Note also that the centre of mass of the satellite varies during the life of the satellite, notably because of the gradual consumption of the onboard fuel. In the known systems, algorithms are implemented for the combined management of the attitude control and of the orbit control, and to make it possible to take account of the position of the centre of mass CM throughout the life of the satellite.
The issue of being able to have effective propulsion systems will therefore be understood. The current solutions, which implement thrusters of different kinds of different locations on the satellite are relatively complex, costly and exhibit a high weight which limits the onboard capacity of the satellite.